The present invention relates to a servo-flap rotor blade system of a rotary-wing aircraft, and more particularly to a servo-flap rotor blade system with an adjustable pitch/flap coupling.
The rotor systems for rotary-wing aircraft such as helicopters and tilt rotor aircraft are relatively complex systems. The rotor system needs to respond to the pilot's input, but also needs to accommodate forces acting upon the rotor system which may be generally outside of direct pilot control.
In a fully articulated rotor system, each individual rotor blade is attached to the rotor hub such that the rotor blade may pitch, lead/lag and flap. Blade pitch is generally under direct management by the pilot through the flight control system. Lead/lag and flapping motion, however, are generally not under the pilot's direct control as these motions are in response to forces from the constantly changing balance between lift, centrifugal, and inertial forces for which the fully articulated hub provides the necessary articulation mechanisms.
Large flapwise loads may cause pitch changes in the rotor blades such that rotor system response may become relatively sluggish, i.e., counteracts cyclic inputs. The detrimental effects of rotor blade flapping, especially in high speed rotary-wing aircraft may be minimized through a pitch-flap coupling (Delta-3) subsystem. Pitch-flap coupling essentially introduces an aerodynamic spring that increases the effective natural frequency of the flap motion to reduce steady and transient blade flapping.
In a fully articulated servo-flap rotor system, there is no direct control link to the rotor blades such that incorporation of a pitch-flap coupling (Delta-3) subsystem may be relatively complicated.
Accordingly, it is desirable to provide a pitch-flap coupling (Delta-3) subsystem for a fully articulated servo-flap rotor system.